Stator with depressions in gaspath wall adjacent leading edges

Abstract

A fluid machine for an aircraft engine has: first and second walls; a gaspath defined between the first wall and the second wall; a rotor having blades rotatable about the central axis; and a stator having: a row of vanes having airfoils including leading edges, trailing edges, pressure sides and suction sides opposed the pressure sides, and depressions defined in the first wall, the depressions extending from a baseline surface of the first wall away from the second wall, a depression of the depressions located circumferentially between a pressure side of the pressure sides and a suction side of the suction sides, the depression axially overlapping the airfoils and located closer to the suction side than to the pressure side, an upstream end of the depression located closer to a leading edge of the leading edges than to a trailing edge of the trailing edges.

The invention claimed is: 
     
       1. A fluid machine for an aircraft engine comprising:
 a first wall and a second wall circumferentially extending around a central axis; 
 a gaspath defined between the first wall and the second wall; 
 a rotor having blades circumferentially distributed around the central axis and extending across the gaspath, the rotor rotatable about the central axis; and 
 a stator in fluid communication with the rotor and having: 
 a row of vanes extending across the gaspath and circumferentially distributed around the central axis, the vanes having airfoils including leading edges, trailing edges, pressure sides and suction sides opposed the pressure sides, and 
 depressions defined in the first wall, the depressions extending from a baseline surface of the first wall away from the second wall, a depression of the depressions located circumferentially between a pressure side of the pressure sides and a suction side of the suction sides, the depression axially overlapping the airfoils and located closer to the suction side than to the pressure side, an upstream end of the depression located closer to a leading edge of the leading edges than to a trailing edge of the trailing edges, the depression intersecting a throat extending from the leading edge to an adjacent suction side of the suction sides. 
 
     
     
       2. The fluid machine of  claim 1 , wherein a ratio of an axial length (h) of the depression taken along an axial direction relative to the central axis to an axial length (C) of the stator taken along the axial direction from the leading edges to the trailing edges ranges from 0.1 to 0.75. 
     
     
       3. The fluid machine of  claim 1 , wherein a ratio of a thickness (t) of the depression taken along a circumferential direction relative to the central axis to a pitch (p) of the stator extending along the circumferential direction from the leading edge of to an adjacent leading edge of the leading edges ranges from 0.05 to 0.5. 
     
     
       4. The fluid machine of  claim 1 , wherein a ratio of a distance (h 1 ) taken along an axial direction relative to the central axis from the upstream end of the depression to the leading edge to an axial length (C) of the stator taken along the axial direction from the leading edges to the trailing edges ranges from −0.25 to 0. 
     
     
       5. The fluid machine of  claim 1 , wherein a ratio of a distance (h 2 ) taken along an axial direction relative to the central axis from a downstream end of the depression to a trailing edge of the trailing edges to an axial length (C) of the stator taken along the axial direction from the leading edges to the trailing edges ranges from 0.25 to 0.75. 
     
     
       6. The fluid machine of  claim 1 , wherein a ratio of a depth (D) of the depression taken along a radial direction relative to the central axis to a span (S) of the airfoils ranges from 0.05 to 0.1. 
     
     
       7. The fluid machine of  claim 1 , wherein a thickness (t) of the depression taken along a circumferential direction relative to the central axis increases along a flow direction of a flow flowing between the airfoils. 
     
     
       8. The fluid machine of  claim 1 , wherein the depression extends parallel to the suction side. 
     
     
       9. The fluid machine of  claim 1 , wherein the fluid machine is a compressor. 
     
     
       10. An aircraft engine comprising:
 a compressor section having: 
 a first wall and a second wall circumferentially extending around a central axis; 
 a gaspath defined between the first wall and the second wall; 
 a rotor having blades circumferentially distributed around the central axis and extending across the gaspath, the rotor rotatable about the central axis; and 
 a stator in fluid communication with the rotor and having: 
 a row of vanes extending across the gaspath and circumferentially distributed around the central axis, the vanes having airfoils including leading edges, trailing edges, pressure sides and suction sides opposed the pressure sides, and 
 depressions defined in the first wall, the depressions extending from a baseline surface of the first wall away from the second wall, a depression of the depressions located circumferentially between a pressure side of the pressure sides and a suction side of the suction sides, the depression axially overlapping the airfoils and closer to the suction side than to the pressure side, an upstream end of the depression located closer to a leading edge of the leading edges than to a trailing edge of the trailing edges, the depression intersecting a throat extending from the leading edge to an adjacent suction side of the suction sides. 
 
     
     
       11. The aircraft engine of  claim 10 , wherein a ratio of an axial length (h) of the depression taken along an axial direction relative to the central axis to an axial length (C) of the stator taken along the axial direction from the leading edges to the trailing edges ranges from 0.1 to 0.75. 
     
     
       12. The aircraft engine of  claim 11 , wherein a ratio of a thickness (t) of the depression taken along a circumferential direction relative to the central axis to a pitch (p) of the stator extending along the circumferential direction from a leading edge of the leading edges to an adjacent leading edge of the leading edges ranges from 0.05 to 0.5. 
     
     
       13. The aircraft engine of  claim 12 , wherein a ratio of a distance (h 1 ) taken along the axial direction from the upstream end of the depression to the leading edge to the axial length (C) of the stator ranges from −0.25 to 0. 
     
     
       14. The aircraft engine of  claim 13 , wherein a ratio of a distance (h 2 ) taken along the axial direction from a downstream end of the depression to a trailing edge of the trailing edges to the axial length (C) of the stator ranges from 0.25 to 0.75. 
     
     
       15. The aircraft engine of  claim 14 , wherein a ratio of a depth (D) of the depression taken along a radial direction relative to the central axis to a span (S) of the airfoils ranges from 0.05 to 0.1. 
     
     
       16. The aircraft engine of  claim 15 , wherein the thickness of the depression increases along a flow direction of a flow flowing between the airfoils. 
     
     
       17. The aircraft engine of  claim 16 , wherein the depression extends parallel to the suction side.